1. Field of the Invention
The present invention relates generally to spacecraft and, more particularly, to spacecraft attitude control systems.
2. Description of the Related Art
The diagram 20 of FIG. 1A illustrates an exemplary spacecraft 22 that orbits in an orbital plane 24 about the earth 26. The spacecraft has a spacecraft body 28 which carries an antenna system 29 and solar panels 30 that generate power for the spacecraft. Although the spacecraft""s orbital plane 24 may be coplanar with the earth""s equatorial plane 32, it is shown, for generality, as having an inclination 34.
The spacecraft 20 includes an attitude control system that maintains a spacecraft service attitude which facilitates the performance of the intended service (e.g., communication service) for which the spacecraft was designed. The spacecraft attitude control system typically responds to attitude measurements from at least one absolute-attitude sensor (e.g., a star tracker) and attitude rate measurements from at least one inertial-attitude sensor (e.g., a gyroscope).
The inertial-attitude sensors are generally arranged to provide attitude rate signals that correspond to three axes (e.g., roll, pitch and yaw axes) of an orbital reference system. Because loss of attitude control implies loss of service, spacecraft typically carry redundant sets of inertial-attitude sensors (or a system of sensors from which more than one set can be configured). Accordingly, the spacecraft""s service can be maintained by substituting a redundant set of inertial-attitude sensors for a failed initial set. This replacement may also be made for other reasons, (e.g., testing to confirm the condition of the redundant set).
Some spacecraft attitude sensors (e.g., staring earth sensors and sun sensors) have wide fields-of-view and others (e.g., star sensors and precision beacon sensors) have more limited fields-of-view. In particular, star trackers are often used in a xe2x80x9cdirect-match modexe2x80x9d of operation after initial attitude has been attained. In this mode, the positions and magnitudes of sensed stars are compared and identified with the aid of a stored star catalog. Although this mode facilitates fast, simple processing, it limits the range over which stars can be identified.
FIG. 1B illustrates an exemplary attitude control system for the spacecraft 22 of FIG. 1A that employs narrow capture range star trackers (e.g., capture range on the order of 0.2 degrees). A graph 40 includes a plot 41 of attitude error about all axes of the local orbital reference and a graph 42 includes a plot 43 of tracked and identified stars. In the simulation, a redundant set of gyroscopes was substituted for an initial set at a time 44.
Plot 41 shows that attitude error about all axes of the local orbital reference (i.e., roll, pitch and yaw axes) remains very low prior to the time 44 and plot 43 shows that at least 4 stars are identified throughout this time. After the time 44, attitude error increases linearly and when it exceeds an error threshold 45, there is a complete loss of identified stars. This degradation of attitude control would cause temporary or even permanent interruption of service of the spacecraft (22 in FIG. 1A).
The present invention is directed to spacecraft methods and structures that enhance attitude control during gyroscope substitutions. The invention recognizes that the error variances of a substituted set of redundant gyroscopes are initially unknown and will introduce significant errors in attitude estimates. If the capture range of absolute-attitude sensors is not significantly larger than these errors, the attitude control system may drive the sensors out of their capture range which endangers the spacecraft""s service.
In response to this recognition, the invention provides methods and structures that temporarily replace an operational process-noise covariance Q of a Kalman filter with a substantially greater interim process-noise covariance Q. This replacement increases the weight given to the most recent attitude measurements and hastens the reduction of attitude errors and gyroscope bias errors. Because greater weight is placed on the most recent attitude measurements and, hence, less weight on the latest predicted attitudes, the error effect of the uncompensated redundant gyroscopes is reduced and the absolute-attitude sensors are not driven out of their capture range.
In another method embodiment, this replacement is preceded by the temporary replacement of an operational measurement-noise variance R with a substantially larger interim measurement-noise variance R to reduce transients during the gyroscope substitutions.
In another method embodiment, an operational error covariance P is temporarily replaced with an substantially greater interim error covariance P.
Method embodiments are also provided for fixed-gain filters.
The novel features of the invention are set forth with particularity in the appended claims. The invention will be best understood from the following description when read in conjunction with the accompanying drawings.